The present invention relates generally to combustors in gas turbine engines and, in particular, to a fuel injection assembly for a gas turbine engine combustor having mixing tubes which are widely dispersed throughout the main combustor dome region.
It will be appreciated that emissions are a primary concern in the operation of gas turbine engines, particularly with respect to the impact on the ozone layer by nitrous oxides (NOx), carbon monoxide (CO), and hydrocarbons. In the case of supersonic commercial transport aircraft flying at high altitudes, current subsonic aircraft technology is not applicable given the detrimental effects on the stratospheric ozone. Accordingly, new fuel injection and mixing techniques have been and continue to be developed in order to provide ultra-low NOx at all engine operating conditions.
In response to such emissions concerns, a new combustor has been developed and is discussed in a parent application entitled xe2x80x9cMulti-Stage Radial Axial Gas Turbine Engine Combustor,xe2x80x9d which is filed concurrently herewith by the assignee of the present invention, has Ser. No. 09/398,577, and is hereby incorporated by reference. It will be seen therein that a key component found to provide extremely low levels of NOx at moderate to high power conditions for aircraft engines was the use of a series of simple mixing tubes as the main fuel injection source. A related patent application entitled xe2x80x9cFuel Flow Control System,xe2x80x9d owned by the assignee of the present invention and having Ser. No. 09/366,510, describes how a control system determines which mixing tubes are to be supplied with fuel in greater detail and is hereby incorporated by reference.
Still, fuel must be transported from a fuel supply controlled by the system in the ""510 patent application into the mixing tubes disclosed in the combustor of the ""577 patent application. It will be appreciated that the mixing tubes are preferably arranged in a plurality of rows and columns. Because the mixing tubes are widely dispersed throughout the main combustor dome region, significant weight, thermal management and structural integrity challenges are presented. As is typical for all flight quality engine hardware, the fuel injection assembly must be as light as possible to minimize engine weight. The thermal management challenge for the fuel injection assembly stems from the extensive fuel-wetted surface area thereof immersed within the high temperature compressor discharge environment, which increases the potential for coke residues to form a partial or full blockage in the fuel passages.
Naturally, the injector tips of the fuel injection assembly must be accurately maintained in position throughout all engine power settings to obtain acceptable system emissions performance. Because the injection sites are widely dispersed, however, maintaining structural integrity of the fuel injection assembly in the hostile dynamic environment of the compressor discharge region, which contains high intensity broadband acoustic excitation, is a particular challenge. Thus, the fuel injection assembly must incorporate sufficient rigidity and damping capability to survive and function in the lightest weight configuration possible.
In light of the foregoing, it would be desirable for a fuel injection assembly to be developed which can provide fuel to a plurality of mixing tubes which are widely dispersed in a gas turbine engine combustor. It would also be desirable for such fuel injection assembly to include continuous active cooling for the fuel stem and injector tip whether fuel is injected into such mixing tubes or not. Further, it would be desirable for the fuel injection assembly to reflect a concern for weight, airflow blockage to the combustor dome region, and ease of removal for maintenance.
In an exemplary embodiment of the invention, a fuel injection assembly for a gas turbine engine combustor is disclosed as including at least one fuel stem, a plurality of concentrically disposed tubes positioned within each fuel stem, wherein a cooling supply flow passage, a cooling return flow passage, and a tip fuel flow passage are defined thereby, and at least one fuel tip assembly connected to each fuel stem so as to be in flow communication with the flow passages, wherein an active cooling circuit for each fuel stem and fuel tip assembly is maintained by providing all active fuel through the cooling supply flow passage and the cooling return flow passage during each stage of combustor operation. The fuel flowing through the active cooling circuit is then collected so that a predetermined portion thereof is provided to the tip fuel flow passage for injection by the fuel tip assembly.